The disclosed embodiments generally pertain to gas turbine engines. More particularly, but not by way of limitation, present embodiments relate to shrouds utilized in gas turbine engines.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially there between. An air inlet or intake is located at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, and a turbine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and low-pressure and high-pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy.
Each of the turbines may include one or more stages of rotor blades which extend radially outward from rotor discs. A shroud assembly circumscribes the turbine rotor and defines an outer boundary for combustion gases flowing through the turbine. The turbine shroud may be a single unitary structure or may be formed of a plurality of segments.
Some known shroud assemblies include a shroud hanger that is coupled to an outer casing of the engine to provide support to a plurality of shrouds positioned adjacent to, and radially outward of, the tips of the turbine blades. It is desirable to limit the clearance between the shroud and the blade tips to limit loss of energy from the combustion gas flowing there through.
The shroud must be capable of meeting the design life requirements for use in the turbine engine operating temperature and pressure environment. To enable current materials to operate effectively in such strenuous temperature and pressure conditions, it has been practiced to utilize composite and, in particular, ceramic matrix composite (CMC) materials for use in the shroud segments because they have higher temperature capability than metallic type parts. However, such ceramic matrix composite (CMC) have mechanical properties that must be considered during the design and application of the CMC use as a shroud segment or component. CMC materials have relatively low tensile ductility or low strain to failure when compared to metallic materials. Also, CMC materials have a coefficient of thermal expansion which differs significantly from metal alloys used as restraining supports or hangers for shrouds of CMC type materials. Therefore, if a CMC shroud segment is restrained and cooled on one surface during operation, forces can develop leading to failure of the segment.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies with gas turbine engines components.